Gas turbine engine component with cusped, lobed cooling hole

ABSTRACT

A component for a gas turbine engine includes a wall and a cooling hole. The wall has a first surface and a second surface. The second surface is exposed to hot gas flow. The cooling hole extends through the wall. The cooling hole includes a metering section extending from an inlet in the first surface of the wall to a transition, a diffusing section extending from the transition to an outlet in the second surface of the wall, a cusp on the transition, and a first longitudinal ridge extending along the diffusing section between the transition and the outlet. The first longitudinal ridge divides the diffusing section into first and second lobes.

CROSS REFERENCE TO RELATED APPLICATION

This application claims priority to U.S. Provisional Application No.61/599,354, filed on Feb. 15, 2012, and entitled “Gas Turbine EngineComponent with Cusped, Lobed Cooling Hole,” the disclosure of which isincorporated by reference in its entirety.

BACKGROUND

This invention relates generally to turbomachinery, and specifically toturbine flow path components for gas turbine engines. In particular, theinvention relates to cooling techniques for airfoils and other gasturbine engine components exposed to hot working fluid flow, including,but not limited to, rotor blades and stator vane airfoils, endwallsurfaces including platforms, shrouds and compressor and turbinecasings, combustor liners, turbine exhaust assemblies, thrust augmentorsand exhaust nozzles.

Gas turbine engines are rotary-type combustion turbine engines builtaround a power core made up of a compressor, combustor and turbine,arranged in flow series with an upstream inlet and downstream exhaust.The compressor section compresses air from the inlet, which is mixedwith fuel in the combustor and ignited to generate hot combustion gas.The turbine section extracts energy from the expanding combustion gas,and drives the compressor section via a common shaft. Expandedcombustion products are exhausted downstream, and energy is delivered inthe form of rotational energy in the shaft, reactive thrust from theexhaust, or both.

Gas turbine engines provide efficient, reliable power for a wide rangeof applications in aviation, transportation and industrial powergeneration. Small-scale gas turbine engines typically utilize aone-spool design, with co-rotating compressor and turbine sections.Larger-scale combustion turbines including jet engines and industrialgas turbines (IGTs) are generally arranged into a number of coaxiallynested spools. The spools operate at different pressures, temperaturesand spool speeds, and may rotate in different directions.

Individual compressor and turbine sections in each spool may also besubdivided into a number of stages, formed of alternating rows of rotorblade and stator vane airfoils. The airfoils are shaped to turn,accelerate and compress the working fluid flow, or to generate lift forconversion to rotational energy in the turbine.

Industrial gas turbines often utilize complex nested spoolconfigurations, and deliver power via an output shaft coupled to anelectrical generator or other load, typically using an external gearbox.In combined cycle gas turbines (CCGTs), a steam turbine or othersecondary system is used to extract additional energy from the exhaust,improving thermodynamic efficiency. Gas turbine engines are also used inmarine and land-based applications, including naval vessels, trains andarmored vehicles, and in smaller-scale applications such as auxiliarypower units.

Aviation applications include turbojet, turbofan, turboprop andturboshaft engine designs. In turbojet engines, thrust is generatedprimarily from the exhaust. Modern fixed-wing aircraft generally employturbofan and turboprop configurations, in which the low pressure spoolis coupled to a propulsion fan or propeller. Turboshaft engines areemployed on rotary-wing aircraft, including helicopters, typically usinga reduction gearbox to control blade speed. Unducted (open rotor)turbofans and ducted propeller engines also known, in a variety ofsingle-rotor and contra-rotating designs with both forward and aftmounting configurations.

Aviation turbines generally utilize two and three-spool configurations,with a corresponding number of coaxially rotating turbine and compressorsections. In two-spool designs, the high pressure turbine drives a highpressure compressor, forming the high pressure spool or high spool. Thelow-pressure turbine drives the low spool and fan section, or a shaftfor a rotor or propeller. In three-spool engines, there is also anintermediate pressure spool. Aviation turbines are also used to powerauxiliary devices including electrical generators, hydraulic pumps andelements of the environmental control system, for example using bleedair from the compressor or via an accessory gearbox.

Additional turbine engine applications and turbine engine types includeintercooled, regenerated or recuperated and variable cycle gas turbineengines, and combinations thereof. In particular, these applicationsinclude intercooled turbine engines, for example with a relativelyhigher pressure ratio, regenerated or recuperated gas turbine engines,for example with a relatively lower pressure ratio or for smaller-scaleapplications, and variable cycle gas turbine engines, for example foroperation under a range of flight conditions including subsonic,transonic and supersonic speeds. Combined intercooled andregenerated/recuperated engines are also known, in a variety of spoolconfigurations with traditional and variable cycle modes of operation.

Turbofan engines are commonly divided into high and low bypassconfigurations. High bypass turbofans generate thrust primarily from thefan, which accelerates airflow through a bypass duct oriented around theengine core. This design is common on commercial aircraft andtransports, where noise and fuel efficiency are primary concerns. Thefan rotor may also operate as a first stage compressor, or as apre-compressor stage for the low-pressure compressor or booster module.Variable-area nozzle surfaces can also be deployed to regulate thebypass pressure and improve fan performance, for example during takeoffand landing. Advanced turbofan engines may also utilize a geared fandrive mechanism to provide greater speed control, reducing noise andincreasing engine efficiency, or to increase or decrease specificthrust.

Low bypass turbofans produce proportionally more thrust from the exhaustflow, generating greater specific thrust for use in high-performanceapplications including supersonic jet aircraft. Low bypass turbofanengines may also include variable-area exhaust nozzles and afterburneror augmentor assemblies for flow regulation and short-term thrustenhancement. Specialized high-speed applications include continuouslyafterburning engines and hybrid turbojet/ramjet configurations.

Across these applications, turbine performance depends on the balancebetween higher pressure ratios and core gas path temperatures, whichtend to increase efficiency, and the related effects on service life andreliability due to increased stress and wear. This balance isparticularly relevant to gas turbine engine components in the hotsections of the compressor, combustor, turbine and exhaust sections,where active cooling is required to prevent damage due to high gas pathtemperatures and pressures.

SUMMARY

This invention concerns a gas turbine engine component with a coolinghole. The wall has a first surface and a second surface. The secondsurface is exposed to hot gas flow. The cooling hole extends through thewall. The cooling hole includes a metering section extending from aninlet in the first surface of the wall to a transition, a diffusingsection extending from the transition to an outlet in the second surfaceof the wall, a cusp on the transition, and a first longitudinal ridgeextending along the diffusing section between the transition and theoutlet. The first longitudinal ridge divides the diffusing section intofirst and second lobes.

Another embodiment of the present invention is a gas turbine enginecomponent that includes a flow path wall having a first surface and asecond surface that is exposed to working fluid flow. An inlet is formedin the first surface of the flow path wall. A metering section extendsfrom the inlet to a transition. A diffusing section extends from thetransition to an outlet in the second surface. A cusp is formed on thetransition. A longitudinal ridge extends along the diffusing sectionfrom the transition toward the outlet. The longitudinal ridge separatesthe diffusing section into first and second lobes.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross-sectional view of a gas turbine engine.

FIG. 2A is a perspective view of an airfoil for the gas turbine engine,in a rotor blade configuration.

FIG. 2B is a perspective view of an airfoil for the gas turbine engine,in a stator vane configuration.

FIG. 3A is a cross-sectional view of the gas path wall for a cooled gasturbine engine component, taken in a longitudinal direction.

FIG. 3B is an alternate cross-sectional view of the gas path wall,showing the cooling hole in a lobed outlet configuration.

FIG. 3C is a cross-sectional view of the gas path wall, taken in atransverse direction.

FIG. 4A is a schematic view of the gas path wall, illustrating a cuspedinlet geometry for the cooling hole.

FIG. 4B is a schematic view of the gas path wall, illustrating analternate cusped inlet geometry for the cooling hole.

FIG. 5A is a schematic view of the gas path wall, illustrating atwo-lobe cooling hole geometry.

FIG. 5B is a schematic view of the gas path wall, illustrating athree-lobe cooling hole geometry.

FIG. 6 is a schematic view of the gas path wall, illustrating a buriedridge geometry.

FIG. 7 is a block diagram illustrating a method for forming a coolinghole in a gas turbine engine component.

DETAILED DESCRIPTION

FIG. 1 is a cross-sectional view of gas turbine engine 10. Gas turbineengine (or turbine engine) 10 includes a power core with compressorsection 12, combustor 14 and turbine section 16 arranged in flow seriesbetween upstream inlet 18 and downstream exhaust 20. Compressor section12 and turbine section 16 are arranged into a number of alternatingstages of rotor airfoils (or blades) 22 and stator airfoils (or vanes)24.

In the turbofan configuration of FIG. 1, propulsion fan 26 is positionedin bypass duct 28, which is coaxially oriented about the engine corealong centerline (or turbine axis) C_(L). An open-rotor propulsion stage26 may also provided, with turbine engine 10 operating as a turboprop orunducted turbofan engine. Alternatively, fan rotor 26 and bypass duct 28may be absent, with turbine engine 10 configured as a turbojet orturboshaft engine, or an industrial gas turbine.

For improved service life and reliability, components of gas turbineengine 10 are provided with an improved cooling configuration, asdescribed below. Suitable components for the cooling configurationinclude rotor airfoils 22, stator airfoils 24 and other gas turbineengine components exposed to hot gas flow, including, but not limitedto, platforms, shrouds, casings and other endwall surfaces in hotsections of compressor 12 and turbine 16, and liners, nozzles,afterburners, augmentors and other gas wall components in combustor 14and exhaust section 20.

In the two-spool, high bypass configuration of FIG. 1, compressorsection 12 includes low pressure compressor (LPC) 30 and high pressurecompressor (HPC) 32, and turbine section 16 includes high pressureturbine (HPT) 34 and low pressure turbine (LPT) 36. Low pressurecompressor 30 is rotationally coupled to low pressure turbine 36 via lowpressure (LP) shaft 38, forming the LP spool or low spool. High pressurecompressor 32 is rotationally coupled to high pressure turbine 34 viahigh pressure (HP) shaft 40, forming the HP spool or high spool.

Flow F at inlet 18 divides into primary (core) flow F_(P) and secondary(bypass) flow F_(S) downstream of fan rotor 26. Fan rotor 26 acceleratessecondary flow F_(S) through bypass duct 28, with fan exit guide vanes(FEGVs) 42 to reduce swirl and improve thrust performance. In somedesigns, structural guide vanes (SGVs) 42 are used, providing combinedflow turning and load bearing capabilities.

Primary flow F_(P) is compressed in low pressure compressor 30 and highpressure compressor 32, then mixed with fuel in combustor 14 and ignitedto generate hot combustion gas. The combustion gas expands to providerotational energy in high pressure turbine 34 and low pressure turbine36, driving high pressure compressor 32 and low pressure compressor 30,respectively. Expanded combustion gases exit through exhaust section (orexhaust nozzle) 20, which can be shaped or actuated to regulate theexhaust flow and improve thrust performance.

Low pressure shaft 38 and high pressure shaft 40 are mounted coaxiallyabout centerline C_(L), and rotate at different speeds. Fan rotor (orother propulsion stage) 26 is rotationally coupled to low pressure shaft38. In advanced designs, fan drive gear system 44 is provided foradditional fan speed control, improving thrust performance andefficiency with reduced noise output.

Fan rotor 26 may also function as a first-stage compressor for gasturbine engine 10, and LPC 30 may be configured as an intermediatecompressor or booster. Alternatively, propulsion stage 26 has an openrotor design, or is absent, as described above. Gas turbine engine 10thus encompasses a wide range of different shaft, spool and turbineengine configurations, including one, two and three-spool turboprop and(high or low bypass) turbofan engines, turboshaft engines, turbojetengines, and multi-spool industrial gas turbines.

In each of these applications, turbine efficiency and performance dependon the overall pressure ratio, defined by the total pressure at inlet 18as compared to the exit pressure of compressor section 12, for exampleat the outlet of high pressure compressor 32, entering combustor 14.Higher pressure ratios, however, also result in greater gas pathtemperatures, increasing the cooling loads on rotor airfoils 22, statorairfoils 24 and other components of gas turbine engine 10. To reduceoperating temperatures, increase service life and maintain engineefficiency, these components are provided with improved coolingconfigurations, as described below. Suitable components include, but arenot limited to, cooled gas turbine engine components in compressorsections 30 and 32, combustor 14, turbine sections 34 and 36, andexhaust section 20 of gas turbine engine 10.

FIG. 2A is a perspective view of rotor airfoil (or blade) 22 for gasturbine engine 10, as shown in FIG. 1, or for another turbomachine.Rotor airfoil 22 extends axially from leading edge 51 to trailing edge52, defining pressure surface 53 (front) and suction surface 54 (back)therebetween.

Pressure and suction surfaces 53 and 54 form the major opposing surfacesor walls of airfoil 22, extending axially between leading edge 51 andtrailing edge 52, and radially from root section 55, adjacent innerdiameter (ID) platform 56, to tip section 57, opposite ID platform 56.In some designs, tip section 57 is shrouded.

Cooling holes or outlets 60 are provided on one or more surfaces ofairfoil 22, for example along leading edge 51, trailing edge 52,pressure (or concave) surface 53, or suction (or convex) surface 54, ora combination thereof. Cooling holes or passages 60 may also be providedon the endwall surfaces of airfoil 22, for example along ID platform 56,or on a shroud or engine casing adjacent tip section 57.

FIG. 2B is a perspective view of stator airfoil (or vane) 24 for gasturbine engine 10, as shown in FIG. 1, or for another turbomachine.Stator airfoil 24 extends axially from leading edge 61 to trailing edge62, defining pressure surface 63 (front) and suction surface 64 (back)therebetween. Pressure and suction surfaces 63 and 64 extend from inner(or root) section 65, adjacent ID platform 66, to outer (or tip) section67, adjacent outer diameter (OD) platform 68.

Cooling holes or outlets 60 are provided along one or more surfaces ofairfoil 24, for example leading or trailing edge 61 or 62, pressure(concave) or suction (convex) surface 63 or 64, or a combinationthereof. Cooling holes or passages 60 may also be provided on theendwall surfaces of airfoil 24, for example along ID platform 66 and ODplatform 68.

Rotor airfoils 22 (FIG. 2A) and stator airfoils 24 (FIG. 2B) are formedof high strength, heat resistant materials such as high temperaturealloys and superalloys, and are provided with thermal anderosion-resistant coatings. Airfoils 22 and 24 are also provided withinternal cooling passages and cooling holes 60 to reduce thermal fatigueand wear, and to prevent melting when exposed to hot gas flow in thehigher temperature regions of a gas turbine engine or otherturbomachine. Cooling holes 60 deliver cooling fluid (e.g., steam or airfrom a compressor) through the outer walls and platform structures ofairfoils 22 and 24, creating a thin layer (or film) of cooling fluid toprotect the outer (gas path) surfaces from high temperature flow.

While surface cooling extends service life and increases reliability,injecting cooling fluid into the gas path also reduces engineefficiency, and the cost in efficiency increases with the requiredcooling flow. Cooling holes 60 are thus provided with improved meteringand inlet geometry to reduce jets and blow off, and improved diffusionand exit geometry to reduce flow separation and corner effects. Coolingholes 60 reduce flow requirements and improve the spread of coolingfluid across the hot surfaces of airfoils 22 and 24, and other gasturbine engine components, so that less flow is needed for cooling andefficiency is maintained or increased.

FIG. 3A is a cross-sectional view of gas turbine engine component(turbine or turbomachinery component) 100 with gas path wall 102, takenin a longitudinal direction and that carries a cool first surface 106and an opposite, hot, second surface 108. Cooling hole 104 extendsthrough gas path wall 102 from first surface 106 to second surface 108to form cooling hole 60 in the, for example outer wall of an airfoil,casing, combustor liner, exhaust nozzle or other gas turbine enginecomponent, as described above.

Gas path wall 102 of component 100 is exposed to cooling fluid on firstsurface 106, with longitudinal hot gas or working fluid flow H alongsecond surface 108. In some components, for example airfoils, firstsurface 106 is an inner surface (or inner wall) and second surface 108is an outer surface (or outer wall). In other components, for examplecombustor liners and exhaust nozzles, first surface 106 is an outersurface (or outer wall), and second surface 108 is an inner surface (orinner wall). More generally, the terms inner and outer are merelyrepresentative, and may be interchanged.

Cooling hole 104 delivers cooling fluid C from first surface 106 of wall102 to second surface 108, for example to provide diffusive flow andfilm cooling. Cooling hole 104 is also inclined along axis A in adownstream direction, in order to improve cooling fluid coverage oversecond surface 108, with less separation and reduced flow mixing.

Cooling hole 104 includes metering section 110 and diffusing section112, and extends along axis A from metering section 110 to diffusingsection 112. Metering section 110 has inlet 114 at first surface 106 ofgas path wall 102, and diffusing section 112 has outlet 116 at secondsurface 108 of gas path wall 102. Outlet 116 defines a perimeter ofdiffusing section 112 at an intersection of diffusing section 112 andsecond surface 108. Surfaces 120, 122, 130, and 132 of cooling hole 104define cooling hole 104 between inlet 114 and outlet 116.

Transition 118 is defined in the region between metering section 110 anddiffusing section 112, where cooling hole 104 becomes divergent(increasing flow area), and where the cooling fluid flow becomesdiffusive. Transition 118 may be relatively abrupt, or may encompass anextended portion of cooling hole 104, for example in a flow transitionregion between metering section 110 and diffusing section 112, or over aregion of overlap between metering section 110 and diffusing section112.

As shown in FIG. 3A, metering section 110 of cooling hole 104 hassubstantially constant or decreasing cross-sectional area in thelongitudinal direction, with upstream and downstream surfaces 120 and122 converging or extending generally parallel to one another along axisA. This maintains or decreases the longitudinal dimension (along thedirection of hot gas flow H) of cooling hole 104, from inlet 114 throughmetering section 110 to transition 118, in order to regulate the coolingfluid flow through inlet 114. Though surfaces 120 and 122 arerepresented in the cross-sectional view of FIG. 3A with a line, they canbe curved as described further below. In the illustrated embodiment,surfaces 120 and 122 are angled with respect to both first surface 106and second surface 108.

Cusp 125 extends longitudinally along metering section 110, from inlet114 toward transition 118. Cusp 125 projects laterally outward (towardaxis A) from downstream surface 122 of cooling hole 104, discouraginglateral flow components to reduce swirl. Cusp 125 also provides meteringcapability, as described below.

Diffusing section 112 of cooling hole 104 diverges between transition118 and outlet 116. That is, upstream and downstream surfaces 120 and122 diverge from one another and away from axis A in the longitudinaldirection, in the region from transition 118 through diffusing section112 to outlet 116. This increases the cross sectional flow area ofdiffusing section 112, in order to provide diffusive flow betweentransition 118 and outlet 116. In the embodiment illustrated in FIG. 3A,diffusing section 112 includes a single lobe 134. Lobe 134 is a surfaceof wall 102 which defines the void of cooling hole 104 at diffusingsection 112. In alternative embodiments, such as FIGS. 5A, 5B, and 6,multiple lobes 134 can be included. In those embodiments, the multiplelobes 134 are surfaces of wall 102 which define distinct channel-likeportions of the void of cooling hole 104 at diffusing section 112.

FIG. 3B is an alternate longitudinal cross-sectional view of gas turbineengine component 100 with gas path wall 102, showing cooling hole 104 ina lobed configuration. In this design, one or more longitudinal ridges124 extend along downstream surface 122 of cooling hole 104, fromtransition 118 toward outlet 116.

Longitudinal ridge 124 projects out (toward axis A) from downstreamsurface 122 of cooling hole 104, discouraging vortex flow and dividingcooling hole 104 into lobes, e.g., in diffusing section 112, in order toreduce swirl and losses at outlet 116. In some designs, diffusingsection 112 can include transition region 128. Transition region 128 canextend from longitudinal ridge 124 to trailing edge 126 of outlet 116,in order to discourage detachment and improve flow uniformity alongsecond surface 108 of gas path wall 102, downstream of cooling hole 104at outlet 116 (see, e.g., FIGS. 5A, 5B, 6). Transition region 128 can beflat or planar. Alternatively, transition region 128 can be non-flat andnon-planar, such as curved (e.g. convex) longitudinally and/or laterallyto encourage flow attachment.

FIG. 3C is a transverse cross sectional view of gas path wall 102, takenalong axis A and looking in a downstream direction, in a planeperpendicular or transverse to the longitudinal cross sections of FIGS.3A and 3B. In this downstream view, hot gas flow H is directed into thepage, and lateral side surfaces 130 and 132 are separated in thetransverse direction across axis A, perpendicular to hot gas flow H.

As shown in the FIG. 3C, metering section 110 of cooling hole 104 hassubstantially constant or decreasing cross sectional flow area. Opposingside surfaces 130 and 132 converge or extend generally parallel to oneanother along axis A in this region, from inlet 114 to transition 118.Thus, metering section 110 acts to restrict or meter cooling fluid flowfrom inlet 114 through transition 118, improving efficiency byregulating the amount of cooling fluid delivered to diffusing section112.

In diffusing section 112, side surfaces 130 and 132 diverge laterallyfrom one another (and from axis A), in the region from transition 118 tooutlet 116. Thus, diffusing section 112 is divergent in both thelongitudinal direction of FIG. 3A, and in the transverse direction ofFIG. 3C. This improves diffusive flow between transition 118 and outlet116, decreasing flow separation at trailing edge 126 and improvingcooling performance along second surface 108 of gas path wall 102. Inthe illustrated embodiment, side surfaces 130 and 132 are substantiallyperpendicular with respect to first surface 106 at their respectiveintersections with first surface 106 and are angled with respect tosecond surface 108 at their respective intersections with second surface108.

FIG. 4A is a schematic view of gas path wall 102, illustrating thecusped geometry of cooling hole 104 in metering section 110. This is adownward or inward view, looking down on second surface 108 of gas pathwall 102, and along cooling hole 104 from outlet 116 toward transition118 and inlet 114. Cusp 125 is presented at the downstream perimeter oftransition 118, extending along metering section 110 of cooling hole 104from transition 118 toward inlet 114. Transition 118 has a substantiallycurved shape that comes to an inwardly-projecting point at cusp 125.Transition 118 can be described as approximately oval shaped, exceptthat cusp 125 creates a pointed projection at the downstream perimeterof inlet 114. Similarly, metering section 110 is approximatelycylindroid-shaped (e.g. an elliptic cylinder), except that cusp 125creates an inward-projecting ridge at a downstream side of meteringsection 110.

Second surface 108 of gas path wall 102 is exposed to hot gas flow H ina longitudinal and downstream direction, from left to right in FIG. 4A.Cooling hole 104 extends down through gas path wall 102, from outlet 116at second surface 108 (solid lines) through transition 118 to inlet 114at first surface 106 (dashed lines). Conversely, metering section 110 ofcooling hole 104 extends upward from inlet 114 to transition 118, anddiffusing section 112 extends upward from transition 118 to outlet 116.

In the particular configuration of FIG. 4A, cusp 125 extends fromtransition 118 to a termination point at inlet 114. The size, length andother geometric properties of cusp 125 are selected to discourage swirl(vortex) flow in cooling hole 104, for example by introducing acanceling vortex pair into the cooling fluid to weaken kidney-shapedvortices formed by crossflow in metering section 110. Cusp 125 may alsobe sized to increase or decrease the cooling flow rate through inlet114, in order to improve flow metering for better coverage and coolingefficiency along second surface 108 of gas path wall 102.

Diffusing section 112 of cooling hole 104 diverges (widens) in both thelongitudinal and lateral directions from transition 118 to outlet 116.This configuration promotes diffusive flow through cooling hole 104,from transition 118 through diffusing section 112 to outlet 116, formore uniform coverage with less detachment along second surface 108 ofgas path wall 102.

The configuration of outlet 116 is also selected to improve coolingperformance. As shown in FIG. 4A, for example, outlet 116 is formed as adelta, with arcuate upstream surface 120 extending toward substantiallylinear trailing edge 126, transverse or perpendicular to hot gas flow H,in order to reduce separation along second surface 108 of gas path wall102. Cusp 125 terminates at transition 118, with cusp 125 tapering as itextends from transition 118 to inlet 114. In the illustrated embodiment,diffusing section 112 is defined as a single undivided lobe betweentransition 118 and outlet 116 with a substantially straight trailingedge 126. Alternatively, trailing edge 126 can be convex (see FIG. 4B).

FIG. 4B is a schematic view of gas path wall 102, illustrating analternate cusped geometry for cooling hole 104. In this configuration,cusp 125 extends from inlet 114 through metering section 110 totransition 118 without substantially tapering as in FIG. 4A. Thus, bothinlet 114 and transition 118 have cusped configurations, as defined by across section taken perpendicular to the axis of cooling hole 104 (see,e.g., FIG. 3B). Outlet 116 has a delta geometry, as described above,with arcuate upstream surface 120 extending to convex trailing edge 126.

The geometries of longitudinal ridge 124 and cusp 125 can vary. Forexample, one or both of ridge 124 and cusp 125 may be formed as long,narrow features extending along the wall of cooling hole 104, where twosloping sides of lobes 134 meet, or as a narrow raised band or ribstructure along (any) surface 120, 122, 130 or 132 of cooling hole 104.Ridge 124 and cusp 125 may also be either substantially pointed orrounded where two arcuate lobes 134 meet, or where the direction ofcurvature reverses along a wall of cooling hole 104. Ridge 124 and cusp125 may also be formed as arched or cone-shape features extending alongthe boundary of two adjacent lobes 134.

FIG. 5A is a schematic view of gas path wall 102, illustrating a lobedconfiguration for cooling hole 104 in diffusing section 112. As shown inFIG. 5A, longitudinal ridge 124 extends along cooling hole 104 betweentransition 118 and outlet 116, dividing cooling hole 104 into lobes 134along diffusing section 112. Thus, lobes 134 define distinctchannel-like portions of the void of cooling hole 104 at diffusingsection 112.

Lobes 134 may have arcuate or curved surfaces along downstream surface122 of diffusing section 112, forming longitudinal ridge 124 as a cuspedridge or rib structure, similar to cusp 125 as described above. As shownin FIG. 5A, moreover, longitudinal ridge 124 and cusp 125 may beco-formed, formed congruently or formed as extensions of one another,with longitudinal ridge 124 and cusp 125 having similar geometry alongdownstream surface 122 of cooling hole 104. Alternatively, longitudinalridge 124 may extend along cooling hole 104 independently of cusp 125,for example from an oval or circular transition 118.

Transition region 128 extends laterally between arcuate extensions 136of longitudinal ridges 124, where arcuate extensions 136 are definedalong the boundaries with adjacent lobes 134. As shown in FIG. 5A, forexample, longitudinal ridge 124 splits or bifurcates into two arcuateextensions 136, which extend longitudinally and transversely alongdiffusing section 112 to trailing edge 126 of outlet 116. In thisparticular configuration, cooling hole 104 has a single flat or planartransition region 128, extending along substantially the entire lengthof trailing edge 126 of outlet 116. Alternatively, transition region 128can be non-flat and non-planar, such as curved (e.g. convex)longitudinally and/or laterally.

FIG. 5B is a schematic view of gas path wall 102, illustrating a threelobe configuration for cooling hole 104 in diffusing section 112. Inthis configuration, two longitudinal ridges 124 extend from transition118 toward outlet 116, dividing cooling hole 104 into three lobes 134.

Two transition regions 128 extend from longitudinal ridges 124 totrailing edge 126 of outlet 116, between adjacent lobes 134. The mutualboundaries of transition regions 128 and adjacent lobes 134 are definedalong arcuate extensions 136, as described above. Transition regions 128extend across substantially all of trailing edge 126, eliminating cuspsand other irregularities to provide more uniform flow coverage andbetter cooling performance along second surface 108 of gas path wall102, downstream of outlet 116.

FIG. 6 is a schematic view of gas turbine engine component 100 with gaspath wall 102, illustrating a “flushed” ridge configuration for coolinghole 104. In this design, longitudinal ridges 124 extend from transition118 toward outlet 116, dividing diffusing section 112 of cooling hole104 into two lobes. However, longitudinal ridge 124 is smoothed out andterminates at transition region 128, as bounded between intersections142 with adjacent outer lobes 134.

Unlike arcuate extensions 136 of longitudinal ridges 124, intersections142 do not extend above downstream surface 122 toward axis A of coolinghole 104. Instead, transition region 128 is defined along downstreamsurface 122, and adjacent lobes 134 curve up from intersections 142toward second (upper) surface 108 of gas path wall 102. Transitionregion 128 extends across substantially all of trailing edge 126,eliminating cusps and other irregularities for more uniform flow.

The overall geometry of cooling hole 104 thus varies, as describedabove, and as shown in the figures. The design of inlet 114 and outlet116 may also vary, including various circular, oblate, oval,trapezoidal, triangular, cusped and delta shaped profiles with arcuateor piecewise linear upstream surfaces 120 and straight or convextrailing edges 126. The configuration of cooling hole 104 is not limitedto these particular examples, moreover, but also encompasses differentcombinations of the various features that are shown, including meteringsections 110 with a variety of different cusps 125, transitions 118 withdifferent circular, elliptical, oblong and cusped cross sections, anddiffusing sections 112 with one, two or three lobes 134, in combinationwith different transition regions 128 bordered by various arcuateextensions 136 and intersections 142. Cusp 125 has point 143 which canbe sharp or rounded. Sides 144 of cusp 125 can be concave, convex, orflat. Moreover, cusp 125 can be asymmetrical, with one of sides 144being different from the other of sides 144. For example, one of sides144 can be concave and the other of sides 144 can be flat.

FIG. 7 is a block diagram illustrating method 200 for forming a coolingflow passage through the gas path wall of a gas turbine enginecomponent. For example, method 200 may be used to form cooling hole 60or cooling hole 104 in an airfoil, casing, liner, combustor, augmentoror turbine exhaust component, as described above.

Method 200 includes forming a cooling hole in a gas path wall of thecomponent (step 202), for example by forming an inlet (step 204),forming a transition (step 206) and forming an outlet (step 208). Method200 may also include forming a metering section (step 210) extendingfrom the inlet to the transition, forming a diffusing section (step 212)extending from the transition to the outlet, and forming ridges todivide the diffusing section into lobes (step 214).

Forming an inlet (step 204) may include forming a cusp on the inlet,where the cusp extends along the metering section from the inlet towardthe transition. The cusp may terminate at the transition, or extend intothe diffusing section, for example extending congruently with alongitudinal ridge.

Forming the diffusing section (step 212) may include forming one or moreridges and lobes (step 214). The lobes are defined by the longitudinalridge, for example extending from the transition toward the outlet.Where a cusp extends to the transition, the longitudinal ridge may beformed as an extension of the cusp, or the cusp may be formed as anextension of the longitudinal ridge.

The gas turbine engine components, gas path walls and cooling holesdescribed herein can thus be manufactured using one or more of a varietyof different processes. These techniques provide each cooling hole withits own particular configuration and features, including, but notlimited to, inlet, metering, transition, diffusion, outlet, upstreamsurface, downstream surface, lateral surface, longitudinal, lobe anddownstream edge features, as described above. In some cases, multipletechniques can be combined to improve overall cooling performance orreproducibility, or to reduce manufacturing costs.

Suitable manufacturing techniques for forming the cooling configurationsdescribed here include, but are not limited to, electrical dischargemachining (EDM), laser drilling, laser machining, electrical chemicalmachining (ECM), water jet machining, casting, conventional machiningand combinations thereof. Electrical discharge machining includes bothmachining using a shaped electrode as well as multiple pass methodsusing a hollow spindle or similar electrode component. Laser machiningmethods include, but are not limited to, material removal by ablation,trepanning and percussion laser machining. Conventional machiningmethods include, but are not limited to, milling, drilling and grinding.

The gas flow path walls and outer surfaces of some gas turbine enginecomponents include one or more coatings, such as bond coats, thermalbarrier coatings, abrasive coatings, abradable coatings and erosion orerosion-resistant coatings. For components having a coating, the inlet,metering section, transition, diffusing section and outlet coolingfeatures may be formed prior to a coating application, after a firstcoating (e.g., a bond coat) is applied, or after a second or third(e.g., interlayer) coating process, or a final coating (e.g.,environmental or thermal barrier) process. Depending on component type,cooling hole or passage location, repair requirements and otherconsiderations, the diffusing section and outlet features may be locatedwithin a wall or substrate, within a thermal barrier coating or othercoating layer applied to a wall or substrate, or combinations thereof.The cooling geometry and other features may remain as described above,regardless of position relative to the wall and coating materials orairfoil materials.

In addition, the order in which cooling features are formed and coatingsare applied may affect selection of manufacturing techniques, includingtechniques used in forming the inlet, metering section, transition,outlet, diffusing section and other cooling features. For example, whena thermal barrier coat or other coating is applied to the second surfaceof a gas path wall before the cooling hole or passage is produced, laserablation or laser drilling may be used. Alternatively, either laserdrilling or water jet machining may be used on a surface without athermal barrier coat. Additionally, different machining methods may bemore or less suitable for forming different features of the coolinghole, for example different laser and other machining techniques may beused for forming the outlet and diffusion features, and for forming thetransition, metering and inlet features.

While the invention is described with reference to exemplaryembodiments, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted withoutdeparting from the spirit and scope of the invention. In addition,different modifications may be made to adapt the teachings of theinvention to particular situations or materials, without departing fromthe essential scope thereof. The invention is thus not limited to theparticular examples disclosed herein, but includes all embodimentsfalling within the scope of the appended claims.

DISCUSSION OF POSSIBLE EMBODIMENTS

The following are non-exclusive descriptions of possible embodiments ofthe present invention.

A component for a gas turbine engine can include a wall and a coolinghole. The wall can have a first surface and a second surface. The secondsurface can be exposed to hot gas flow. The cooling hole can extendthrough the wall. The cooling hole can include a metering sectionextending from an inlet in the first surface of the wall to atransition, a diffusing section extending from the transition to anoutlet in the second surface of the wall, a cusp on the transition, anda first longitudinal ridge extending along the diffusing section betweenthe transition and the outlet. The first longitudinal ridge can dividethe diffusing section into first and second lobes.

The component of the preceding paragraph can optionally include,additionally and/or alternatively any, one or more of the followingfeatures, configurations and/or additional components:

the cooling hole can extend through the wall along an axis, and thefirst longitudinal ridge can extend away from a downstream surface ofthe cooling hole toward the axis;

the cusp can extend from the transition along the metering section ofthe cooling hole toward the inlet;

the cusp can extend to the transition and the longitudinal ridge canextend from the cusp at the transition along the diffusing sectiontoward the outlet;

a third lobe can be positioned between the first and second lobes, thefirst longitudinal ridge can divide the first lobe from the third lobe;and a second longitudinal ridge can divide the second lobe from thethird lobe;

the third lobe can terminate in a transition region extending from thefirst and second longitudinal ridges to a trailing edge of the outlet;

a first transition region can extend from the first longitudinal ridgeto a trailing edge of the outlet and a second transition region canextend from the second longitudinal ridge to the trailing edge of theoutlet;

the outlet can include a trailing edge that is substantially straightalong the second surface of the wall;

the outlet can include a trailing edge that is substantially convex;

the cusp can taper as it extends from the transition to the inlet;and/or

the wall can form one of a pressure surface, a suction surface or aplatform surface of an airfoil.

A gas turbine engine component can include a flow path wall having afirst surface and a second surface that is exposed to working fluidflow. An inlet can be formed in the first surface of the flow path wall.A metering section can extend from the inlet to a transition. Adiffusing section can extend from the transition to an outlet in thesecond surface. A cusp can be formed on the transition. A longitudinalridge can extend along the diffusing section from the transition towardthe outlet. The longitudinal ridge can separate the diffusing sectioninto first and second lobes.

The gas turbine engine component of the preceding paragraph canoptionally include, additionally and/or alternatively any, one or moreof the following features, configurations and/or additional components:

the cusp can extend from the transition along the metering sectiontoward the inlet;

the first longitudinal ridge can be a single longitudinal ridgeextending from the cusp at the transition and the first longitudinalridge can extend toward the outlet to divide the diffusing section intothe first and second lobes;

a third lobe can be positioned between the first and second lobes, thefirst longitudinal ridge can divide the first lobe from the third lobe,and a second longitudinal ridge can divide the second lobe from thethird lobe;

the third lobe can terminate in a transition region extending from thefirst and second longitudinal ridges to a trailing edge of the outlet;

the cusp can taper as it extends from the transition to the inlet;

the outlet can include a trailing edge that is substantially straightalong the second surface of the wall;

the outlet can include a trailing edge that is substantially convex;and/or

a turbofan engine can include the gas turbine engine component.

1. A component for a gas turbine engine, the component comprising: a wall having a first surface and a second surface, wherein the second surface is exposed to hot gas flow; and a cooling hole extending through the wall, the cooling hole comprising: a metering section extending from an inlet in the first surface of the wall to a transition; a diffusing section extending from the transition to an outlet in the second surface of the wall; a cusp on the transition; and a first longitudinal ridge extending along the diffusing section between the transition and the outlet, wherein the first longitudinal ridge divides the diffusing section into first and second lobes.
 2. The component of claim 1, wherein the cooling hole extends through the wall along an axis, and wherein the first longitudinal ridge extends away from a downstream surface of the cooling hole toward the axis.
 3. The component of claim 1, wherein the cusp extends from the transition along the metering section of the cooling hole toward the inlet.
 4. The component of claim 3, wherein the cusp extends to the transition and the longitudinal ridge extends from the cusp at the transition along the diffusing section toward the outlet.
 5. The component of claim 4, and further comprising: a third lobe positioned between the first and second lobes, wherein the first longitudinal ridge divides the first lobe from the third lobe; and a second longitudinal ridge dividing the second lobe from the third lobe.
 6. The component of claim 5, wherein the third lobe terminates in a transition region extending from the first and second longitudinal ridges to a trailing edge of the outlet.
 7. The component of claim 5, further comprising: a first transition region extending from the first longitudinal ridge to a trailing edge of the outlet; and a second transition region extending from the second longitudinal ridge to the trailing edge of the outlet.
 8. The component of claim 1, wherein the outlet comprises: a trailing edge that is substantially straight along the second surface of the wall.
 9. The component of claim 1, wherein the outlet comprises: a trailing edge that is substantially convex.
 10. The component of claim 1, wherein the cusp tapers as it extends from the transition to the inlet.
 11. The component of claim 1, wherein the wall forms one of a pressure surface, a suction surface or a platform surface of an airfoil.
 12. A gas turbine engine component comprising: a flow path wall having a first surface and a second surface, wherein the second surface is exposed to working fluid flow; an inlet formed in the first surface of the flow path wall; a metering section extending from the inlet to a transition; a diffusing section extending from the transition to an outlet in the second surface; a cusp formed on the transition; and a first longitudinal ridge extending along the diffusing section from the transition toward the outlet, wherein the first longitudinal ridge separates the diffusing section into first and second lobes.
 13. The gas turbine engine component of claim 12, wherein the cusp extends from the transition along the metering section toward the inlet.
 14. The gas turbine engine component of claim 13, wherein the first longitudinal ridge is a single longitudinal ridge extending from the cusp at the transition, the first longitudinal ridge extending toward the outlet to divide the diffusing section into the first and second lobes.
 15. The gas turbine engine component of claim 13, and further comprising: a third lobe positioned between the first and second lobes, wherein the first longitudinal ridge divides the first lobe from the third lobe; and a second longitudinal ridge dividing the second lobe from the third lobe.
 16. The gas turbine engine component of claim 15, wherein the third lobe terminates in a transition region extending from the first and second longitudinal ridges to a trailing edge of the outlet.
 17. The gas turbine engine component of claim 13, wherein the cusp tapers as it extends from the transition to the inlet.
 18. The gas turbine engine component of claim 12, wherein the outlet comprises: a trailing edge that is substantially straight along the second surface of the wall.
 19. The gas turbine engine component of claim 12, wherein the outlet comprises: a trailing edge that is substantially convex.
 20. A turbofan engine comprising the gas turbine engine component of claim
 12. 